Space propulsion systems vary widely in both design and performance. Performance, as measured by specific impulse, determines the ratio of usable payload to propellant mass required to propel the system to its destination. High specific impulses allow greater usable payload masses, thereby allowing greater payloads of existing missions and new space missions that otherwise could not be achieved, and at potentially reduced costs.
Solar thermal propulsion systems have been proposed as means to achieve greater payload fractions. These engines can be used, for example, to boost payloads from low earth orbits to higher orbits. In such an engine, solar radiation is captured and focused by mirrors into a "black body" cavity of the unit, where the solar radiation heats a propellant, such as hydrogen. The propellant is then passed through a nozzle, creating thrust.
An early solar thermal propulsion engine, described by Sanders in U.S. Pat. No. 3,064,418, contains a pebble bed heat exchanger. The sun's rays are admitted through windows to a heat exchanger containing a pebble bed of refractory material. Propellant is heated as it passes through the heat exchanger and is passed through a converging/diverging rocket nozzle, creating thrust. Coolant propellant is passed through the chamber walls, so as to cool them below their material temperature limits and recover thermal energy to the propellant that might otherwise be lost. This solar thermal propulsion engine is complex to build and difficult to operate. Its internal cavity geometry promotes high re-radiation loss out its cavity entrance, hence giving poor thermal efficiency.
One solar thermal propulsion device uses one or more series of coiled refractory hollow metal tubes configured to form a conical or cylindrical shaped solar collection cavity, such as shown for the solar energy focusing assembly and storage unit described by Vrolyk et al in U.S. Pat. No. 4,815,443. Focused solar energy is directed into this cavity and is absorbed by the metal tubes. Hydrogen gas passing through the tubes is heated to high temperature. High temperature gas is then directed to a rocket nozzle where it is expended out of the nozzle, creating thrust. This device is complex to build and has low thermal efficiency due to the large re-radiation losses out of the cavity entrance.
Another solar thermal propulsion device, described by Shoji in "Potential of Advanced Solar Thermal Propulsion", Orbit-Raising and Maneuvering Propulsion: Research Status and Needs, Volume 89, American Institute of Aeronautics and Astronautics, Inc. (Ed. L. H. Caveny, 1984), consists of a deep solar collection cavity, equipped with an optically clear window at its entrance. Focused solar energy passes through the window and into the cavity. Hydrogen mixed with metal alkali seed particles are injected into the cavity, where solar energy is first absorbed by the seed particles and, in turn, is transferred to the hydrogen by physical contact. Heated hydrogen is then expended out of a rocket nozzle. This device has the added complexity of an optical window and separate seed particle feed system, and low thermal efficiency because the window is opaque to some incoming frequencies of sunlight. It also has large re-radiation losses out of the cavity entrance. Furthermore, when operating at high specific impulses (low hydrogen flow rates), there is inadequate hydrogen flow to actively cool the window.
An additional solar thermal propulsion device, described by Shoji in U.S. Pat. No. 4,781,018, uses the above mentioned windowed cavity with the addition of a series of porous disks. Solar energy passing through the window impinges on an optically coarse porous disk. The disk absorbs a fraction of the solor energy and passes the remaining solar energy to underlying disks located deeper within the cavity. Hydrogen enters the cavity by a series of jet vanes directed at the window for purposes of actively cooling the window. Hydrogen then flows through the porous disks, absorbing solar energy by contact and in turn cooling the disks, minimizing re-radiation losses back out the window. The porous disk concept is specific impulse-limited by the hydrogen flow requirements of actively cooling the window. Re-radiation losses through the window also increase when operating at high specific impulses, because the first porous disk cannot be adequately cooled by the resulting low hydrogen mass flow rates.
The inventor has studies the field of solar thermal propulsion engines and developed a number of novel solar thermal engine designs. One objective is to absorb solar energy efficiently, transferring it to the propellant and expending it out of a nozzle to create thrust. Another objective is to achieve high specific-impulse, on the order of 800 to 900 or more pounds-force-second/pounds-mass (1bf-sec/1bm), by maximizing final propellant temperature. Yet another objective is to have a compact heat exchanger by way of high transfer rates per unit area, promoting high efficiency and low weight. A further objective is towards simplicity, reliability, and safety through use of low number part count, state-of-the-art fabrication techniques and well characterized materials.